Turbine rotor blade of a gas turbine and method for cooling a blade tip of a turbine rotor blade of a gas turbine

ABSTRACT

The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, on which means for duct-type guidance of cooling air extending from a front suction-side area of the blade tip to a rear area of the blade tip are provided, and to a method for cooling a blade tip of a turbine rotor blade of a gas turbine, where air from a hot gas flow is guided from a front suction-side area of a blade tip to a rear area of the blade tip through a duct-type guidance.

This application claims priority to German Patent ApplicationDE102013224998.5 filed Dec. 5. 2013, the entirety of which isincorporated by reference herein.

This invention relates to a turbine rotor blade of a gas turbine with ablade tip. Furthermore, this invention relates to a method for coolingsuch a blade tip of a turbine rotor blade.

It is known from the state of the art that a leakage mass flow caused bythe pressure difference from a blade pressure side to a blade suctionside arises at a radial gap between a turbine rotor and a casing.Attempts are therefore being made to design the blade tip of the turbinerotor such that the leakage mass flow is reduced. Another objective isto reduce the negative effect of the blade tip leakage vortex caused bythe leakage mass flow on the turbine aerodynamics.

To improve the flow over the blade tips of the turbine rotor,circumferential sealing edges (squealers) are used. Designs are alsoknown, where overhangs at the blade tip (winglets) are provided. Thecircumferential sealing edges can contribute to an improvement in theaerodynamics. The overhangs on the suction side and/or on the pressureside can reduce the leakage mass flow and also improve the aerodynamicsaround the blade tip.

It is furthermore necessary in turbine rotor blades, particularly underoperating conditions with high thermal loads due to the hot gas flow, toachieve high durability and a long service life for the blade tips. Tothat end, it is known to cool the turbine rotor blades internally byconvection using compressor air and to discharge this air for thepurpose of film cooling on the blade outer surfaces or blade tips.Besides the blade leading edge, the thermally most highly loaded area ofa blade tip is often the rear pressure-side area. It is known to provideon the blade tip pressure-side holes or recesses through whichblade-internal film cooling air is passed out near the blade tip. Inaddition, cooling air can be passed out through openings on the bladetip (dust holes), in order to prevent dirt accumulations in the internalcooling air passages and to achieve additional cooling.

The measures known from the state of the art entail a number ofdisadvantages:

When film cooling air is used, it is necessary to consume a largequantity of “expensive” film cooling air in order to protect the bladetip from the high temperatures of the hot gas. The film cooling air mustbe taken from the compressor of the gas turbine, thereby reducing theefficiency of the thermodynamic cyclic process inside the gas-turbineengine.

Convective internal cooling of the blade tip by internal cooling airpassages is relatively ineffective in the area of the blade tip. This ispartly due to the fact that only insufficiently high internal heattransfer coefficients can be generated directly underneath the bladetip. In addition, conventional casting techniques, using which theturbine rotor blades are manufactured, require relatively high minimumwall thicknesses, so that the temperature in the outer wall area isrelatively close to the hot gas temperature.

Cooling air that must be forcibly discharged at the blade tip in orderto prevent dirt accumulations in the internal blade ducts contributesonly to a limited extent to the cooling of the blade tip, since itcannot reach the thermally most highly loaded points of the blade tip.

The object underlying the present invention is to provide a method forcooling a blade tip of a turbine rotor blade of a gas turbine as well asa turbine rotor blade suitable for carrying out the method, which, whilebeing simply designed guarantees effective cooling.

It is a particular object of the present invention to provide solutionto the above problematics by a combination of the features of theindependent Claims. Further advantageous embodiments of the inventionbecome apparent from the sub-claims.

In accordance with the invention, it is thus provided in respect of theturbine rotor blade that means for duct-type guidance of cooling airextend from its front suction-side area to a rear area. The inventionthus provides for the thermally highly loaded rear area of the turbinerotor blade tip to be supplied at least partly with passive air, i.e.with relatively cold air from the hot gas flow. It is self-evident thatthe hot gas temperature must be below the maximum temperaturesustainable by the blade material. There is an area with relatively lowhot gas temperature on the front suction-side blade tip. In particularin the case of additional discharge of cooling air at the turbine casingupstream of the rotor blades, the hot gas temperature in this area canbe considerably below the highest permissible metal temperature.

In accordance with the invention, the relatively “cold” hot gas flow isthus used for cooling the blade tip by being routed to the thermallymost highly loaded pressure-side rear area of the blade tip. This isachieved by the means provided in accordance with the invention forduct-type guidance of this cooling air. In a particularly favourableembodiment of the invention, it is provided that a cover forming a flowduct is attached at the blade tip. This creates a duct or a cavity onthe blade tip, through which the cooling air can be passed. This duct orcavity, forming the means in accordance with the invention for duct-typeguidance of cooling air, preferably has an inlet opening on thesuction-side blade leading edge. It is however also possible tointroduce cooling air from the blade interior. In the area of the bladetrailing edge too, an opening is provided through which the cooling airflows out. The pressure difference over the blade row ensures here aflow of relatively cold hot gas through the duct or cavity.

The means provided in accordance with the invention can extend over theentire length or only over part of the length of the blade tip.

The cooling air can exit the duct or cavity in accordance with theinvention at the blade tip, depending on the required counter-pressurelevel, in the area of the pressure-side or the suction-side rear bladetip. The discharge at the pressure side has the further advantage,besides lower aerodynamic losses, that the cooling air discharged thereis in turn sucked into the blade tip gap, so that the external blade tipsurface too is supplied with relatively cold air.

In a preferred embodiment of the invention, a protective cover asmentioned is fitted onto a circumferential sealing edge of the blade tip(winglet, squealer) and fastened there. The result is an effective, flatand contourless blade tip geometry with a duct or cavity underneath. Inan even more advantageous embodiment of the invention, it is providedthat the cover is suitably shaped or contoured such that the blade tipcan retain the contour of the circumferential sealing edge.

It is particularly advantageous when the wall thicknesses of the coverare designed relatively thin, so that its maximum metal temperature canbe kept as low as possible.

The protective cover and/or the cavity on the blade tip can be butdo/does not necessarily have to be, designed up to the blade leadingedge. Since the pressure level in the front blade tip area is relativelyconstant and only drops steeply towards the blade trailing edge, it canbe advantageous to design the protective cover and hence the duct orcavity only starting from a middle position of the blade tip. With thisembodiment, the dust holes can then be near the cover and hence close tothe inflow area of the duct or cavity, in order to promote theaspiration of the cold dust hole air into the duct or cavity. It is thuspossible to use the dust hole air, otherwise not readily usable forcooling, to cool the pressure-side blade tip close to the trailing edge.

In a further embodiment of the invention, it is possible to design theprotective cover at the blade leading edge closed (without opening).Hence only one opening of the cavity or of the duct is provided close tothe blade trailing edge. With this embodiment of the invention, thecavity or duct can be flooded completely with cold blade-internal air.This embodiment is suitable in particular for very high hot gastemperatures, when blade-external air is no longer usable for coolingthe blade tip.

For guiding the flow inside the blade tip cavity or in the flow ductprovided at the blade tip, additional webs or supports can be used. Bymeans of these webs, the air can be guided to the thermally most highlyloaded areas. The webs or supports are furthermore used as fasteningsurfaces for the protective cover and thereby contribute to themechanical stability of the blade tip. Furthermore, the webs or supportscan increase the internal heat transfer in the cavity or duct, so thatthe cooling effect can be further improved.

The following advantages result in accordance with the invention, asalready partially explained above:

With the invention, the quantity of film cooling air required forcooling the rotor blade tip can be considerably reduced.

Due to the more effective cooling of the blade tip and the reduction ofthe blade tip temperatures that this entails, the wear on the blade tipcan be reduced and hence the service life of the turbine blade extended.

Due to the reduced wear on the blade tip during operation, the decreasein turbine efficiency over the period of operation can be reduced.

In accordance with the invention the operating costs of the gas turbineare reduced.

The present invention is described in the following in light of theaccompanying drawing showing exemplary embodiments. In the drawing,

FIG. 1 shows a schematic representation of a gas-turbine engine inaccordance with the present invention,

FIG. 2 shows a simplified sectional view of a blade tip designed inaccordance with the present invention,

FIG. 3 shows a view, by analogy with FIG. 2, of a further exemplaryembodiment of the present invention,

FIGS. 4 shows a further exemplary embodiment, by analogy with FIGS. 2and 3 without lateral blade tip overhang, and

FIGS. 5 to 12 show simplified perspective representations of exemplaryembodiments in accordance with the present invention.

The gas-turbine engine 10 in accordance with FIG. 1 is a generallyrepresented example of a turbomachine where the invention can be used.The engine 10 is of conventional design and includes in the flowdirection, one behind the other, an air inlet 11, a fan 12 rotatinginside a casing, an intermediate-pressure compressor 13, a high-pressurecompressor 14, a combustion chamber 15, a high-pressure turbine 16, anintermediate-pressure turbine 17 and a low-pressure turbine 18 as wellas an exhaust nozzle 19, all of which being arranged about a centralengine axis 1.

The intermediate-pressure compressor 13 and the high-pressure compressor14 each include several stages, of which each has an arrangementextending in the circumferential direction of fixed and stationary guidevanes 20, generally referred to as stator vanes and projecting radiallyinwards from the engine casing 21 in an annular flow duct through thecompressors 13, 14. The compressors furthermore have an arrangement ofcompressor rotor blades 22 which project radially outwards from arotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine16 or the intermediate-pressure turbine 17, respectively.

The turbine sections 16, 17, 18 have similar stages, including anarrangement of fixed stator vanes 23 projecting radially inwards fromthe casing 21 into the annular flow duct through the turbines 16, 17,18, and a subsequent arrangement of turbine rotor blades 24 projectingoutwards from a rotatable hub 27. The compressor drum or compressor disk26 and the blades 22 arranged thereon, as well as the turbine rotor hub27 and the turbine rotor blades 24 arranged thereon rotate about theengine axis 1 during operation.

FIG. 2 shows a simplified sectional view of a blade tip 29 of a turbinerotor blade 24. The reference numeral 35 shows the suction side, whilethe reference numeral 36 indicates the pressure side. Front-side sealingedges 33 are provided on the blade tip 29. Furthermore, the blade tip 29can have lateral overhangs (winglets) 34. A flow duct 37 is provided onthe front side of the blade tip 29 and closed by a protective cover 38.The protective cover 38 is designed profiled, so that any contouring ofthe blade tip 29 can be retained.

FIG. 3 shows a view, by analogy with FIG. 2, where additional webs orsupports 39 are provided which support the protective cover 38. The websor supports 39 furthermore enable several flow ducts 37 to be formed, orthe airflow through the flow duct 37 to be optimized.

FIG. 4 shows a view, by analogy with FIGS. 2 and 3, where in theexemplary embodiment of FIG. 4 the blade tip has no lateral overhang(winglet overhang).

FIGS. 5 to 12 show differing exemplary embodiments of the invention,where the perspective view is schematic and where the protective cover38 is only shown in simplified form in order to make clear the flowthrough the flow duct 37.

In all exemplary embodiments of FIGS. 5 to 12, the front suction-sidearea of the blade tip 29 has the reference numeral 30, while the reararea has the reference numeral 31. A blade trailing edge 32 is designedin the usual way.

FIG. 5 shows an exemplary embodiment in which the protective cover 38 isprovided on the entire front-side area of the blade tip 29. The flownecessary for cooling the blade tip is introduced centrally at the bladeleading edge, while the outflow through a suction-side opening takesplace in the rear area of the blade tip.

FIG. 6 shows a design variant of the exemplary embodiment of FIG. 5,where the protective cover 38 extends over only part of the total lengthof the blade tip 29.

Complementing the design variant of FIG. 5, a centric support 39 isprovided in the exemplary embodiment of FIG. 7, which divides the flowduct 37.

In the exemplary embodiment of FIG. 8, the outflow is provided, in avariation of the exemplary embodiment of FIG. 7 on the pressure side ofthe rear area 31 of the blade tip 29.

The exemplary embodiment of FIG. 9 represents a variant of the exemplaryembodiments of FIGS. 7 and 8 and has a centric outlet opening in thearea of the blade trailing edge 32.

In the exemplary embodiment of FIG. 10, the protective cover extendscompletely over the front area of the blade tip 29 and has an outletopening only on the suction side of the rear area 31. The cooling air issupplied from the blade interior via ducts.

The exemplary embodiment of FIG. 11 also shows a protective cover closedin the front area, by analogy with FIG. 10, where a central web 39divides the flow duct 37.

In the exemplary embodiment of FIG. 12, it is provided, in a variationfrom the exemplary embodiments of FIGS. 10 and 11, that individualsupports acting as turbulators are arranged inside the flow duct 37.

The exemplary embodiments of FIGS. 10 to 12 each show the supply of theflow through cooling air holes (dust holes) 41.

List of Reference Numerals

-   1 Engine axis-   10 Gas-turbine engine/core engine-   11 Air inlet-   12 Fan-   13 Intermediate-pressure compressor (compressor)-   14 High-pressure compressor-   15 Combustion chamber-   16 High-pressure turbine-   17 Intermediate-pressure turbine-   18 Low-pressure turbine-   19 Exhaust nozzle-   20 Guide vanes-   21 Engine casing-   22 Compressor rotor blades-   23 Stator vanes-   24 Turbine rotor blades-   25 —-   26 Compressor drum or disk-   27 Turbine rotor hub-   28 Exhaust cone-   29 Blade tip-   30 Front suction-side area of blade tip 29-   31 Rear area of blade tip 29-   32 Blade trailing edge-   33 Sealing edge-   34 Overhang-   35 Suction side-   36 Pressure side-   37 Flow duct-   38 Cover/protective cover-   39 Web/support-   40 Casing-   41 Cooling air hole

1. Turbine rotor blade of a gas turbine with a blade tip, on which meansfor duct-type guidance of cooling air extending from a frontsuction-side area of the blade tip to a rear area of the blade tip areprovided.
 2. Turbine rotor blade in accordance with claim 1, wherein themeans are provided at least over part of the length of the blade tip. 3.Turbine rotor blade in accordance with claim 1, wherein the means forthe inlet of cooling air from the hot gas flow are provided in the frontsuction-side area of the blade tip.
 4. Turbine rotor blade in accordancewith claim 1, wherein the means for the outlet of cooling air areprovided in the pressure-side area of the rear blade tip or in thesuction-side area of the rear blade tip or in the area of the bladetrailing edge.
 5. Turbine rotor blade in accordance with claim 1,wherein a sealing edge and/or an overhang are/is provided at the bladetip.
 6. Turbine rotor blade in accordance with claim 1, wherein themeans are designed in the form of a cover arranged at the blade tip andforming a flow duct.
 7. Turbine rotor blade in accordance with claim 1,wherein the means are designed for the introduction of air exiting atleast one air duct extending inside the turbine rotor blade.
 8. Turbinerotor blade in accordance with claim 1, wherein at least one web orsupport is provided at the blade tip for guiding the cooling air in theflow duct.
 9. Turbine rotor blade in accordance with claim 6, whereinthe cover is arranged on the web.
 10. Method for cooling a blade tip ofa turbine rotor blade of a gas turbine, where air from a hot gas flow isguided from a front suction-side area of a blade tip to a rear area ofthe blade tip through a duct-type guidance.